1. Field of the Invention
The present invention relates generally to a cooling system and method for a gas turbine engine and, in particular, to a system and method of providing clean cooling air to a hot portion of a gas turbine engine.
2. Description of Related Art
Gas turbine engines (such as turbojet engines, bypass turbofan engines, turboprop engines, turboshaft engines, etc.) may be used to power flight vehicles (such as planes, helicopters, and missiles, etc.) and may also be used to power ships tanks, electric power generators, pipeline pumping apparatus, etc. For purposes of illustration, the present invention will be described with respect to an aircraft bypass turbofan gas turbine engine. However, it is understood that the invention is equally applicable to other types and/or uses of gas turbine engines.
A gas turbine engine includes a core engine having, in serial flow relationship, a high pressure compressor (also called a core compressor) to compress the airflow entering the core engine, a combustor in which a mixture of fuel and the compressed air is burned to generate a propulsive gas flow, and a high pressure turbine which is rotated by the propulsive gas flow and which is connected by a larger diameter shaft to drive the high pressure compressor. A typical aircraft bypass gas turbine engine adds a low pressure turbine (located aft of the high pressure turbine) which is connected by a smaller diameter coaxial shaft to drive a front fan (located forward of the high pressure compressor) which is surrounded by a nacelle and which may also drive a low pressure compressor (located between the front fan and the high pressure compressor). The low pressure compressor sometimes is called a booster compressor or simply a booster. It is understood that the term "compressor" includes, without limitation, high pressure compressors and low pressure compressors. A flow splitter, located between the fan and the first (usually the low pressure) compressor, separates the air which exits the fan into a core engine airflow and a surrounding bypass airflow. The bypass airflow from the fan exits the fan bypass duct to provide most of the engine thrust for the aircraft. Some of the engine thrust comes from the core engine airflow after it flows through the low and high pressure compressors to the combustor and is expanded through the high and low pressure turbines and accelerated out of the exhaust nozzle.
Aircraft bypass turbofan gas turbine engines are designed to operate at high temperatures to maximize engine thrust. Cooling of engine hot section components (such as the combustor, the high pressure turbine, the low pressure turbine, and the like) is necessary because of the thermal "redline" limitations of the materials used in the construction of such components. Typically, such cooling of a portion of the engine is accomplished by ducting (also called "bleeding") cooler air from the high and/or low pressure compressors to those engine components which require such cooling. Unfortunately, the relatively low pressure and hot temperature of the compressor air limits its ability to be used to cool such engine components.
In order to improve cooling of the hot section components and other portions of a gas turbine engine, a gas turbine engine cooling system disclosed by U.S. Pat. No. 5,392,614 was developed which includes a turbocompressor and a heat exchanger. As seen therein, this system provides air having a higher pressure and a lower temperature for cooling portions of the engine, such as hot section components.
One manner of cooling a hot object is by means of transpiration cooling, where cooling air is passed through a porous surface in order to provide a continuous sheet of air insulation on the hot side of the surface. This type of cooling in a gas turbine engine has heretofore been impossible to use because of dirt and other contaminants in the cooling air which would plug the tiny cooling holes (on the order of 1/1000 of an inch versus holes having a range of 10-15/1000 of an inch). To the extent the cooling air might have been filtered, this has been made impractical due to the pressure loss resulting therefrom. This problem of contaminants in the cooling air is also applicable to film cooling, as the holes utilized therefor must remain large enough to permit dirty cooling air to pass therethrough.
Accordingly, it is desired that a system for a gas turbine engine be developed which enables cooling air to be filtered without experiencing a net pressure loss.